Engine bleed system with multi-tap bleed array

ABSTRACT

An engine bleed control system for a gas turbine engine of an aircraft is provided. The engine bleed control system includes a multi-tap bleed array including engine bleed taps coupled to a compressor source of a lower pressure compressor section before a highest pressure compressor section of the gas turbine engine. A highest stage of the engine bleed taps has a maximum bleed temperature below an auto-ignition point of a fuel-air mixture of the aircraft at idle engine power at a maximum aircraft altitude and a pressure suitable for pressurizing the aircraft at the maximum aircraft altitude. The engine bleed control system also includes a plurality of valves operable to extract bleed air from each of the engine bleed taps. A controller is operable to selectively open and close each of the valves based on a bleed air demand and control delivery of the bleed air to an aircraft use.

BACKGROUND

This disclosure relates to gas turbine engines, and more particularly toan engine bleed system with a multi-tap bleed array.

Gas turbine engines are used in numerous applications, one of which isfor providing thrust to an aircraft. Compressed air is typically tappedat a high pressure location near the combustor for auxiliary uses, suchas environmental control of the aircraft. However, this high pressureair is typically hotter than can safely be supported by ductwork anddelivery to the aircraft. Thus, a pre-cooler or heat exchanger is usedto cool high-temperature engine bleed air and is typically located nearthe engine such that excessively hot air is not ducted through the wingof the aircraft for safety reasons. Diverting higher pressure and highertemperature air from the engine beyond the pressure needed reducesengine efficiency. Further, heat exchangers used to cool engine bleedair add to overall aircraft weight, which also reduces fuel burnefficiency.

BRIEF DESCRIPTION

According to an embodiment, an engine bleed control system for a gasturbine engine of an aircraft is provided. The engine bleed controlsystem includes a multi-tap bleed array including a plurality of enginebleed taps coupled to a compressor source of a lower pressure compressorsection before a highest pressure compressor section of the gas turbineengine. A highest stage of the engine bleed taps has a maximum bleedtemperature below an auto-ignition point of a fuel-air mixture of theaircraft at idle engine power at a maximum aircraft altitude and apressure suitable for pressurizing the aircraft at the maximum aircraftaltitude. The engine bleed control system also includes a plurality ofvalves operable to extract bleed air from each of the engine bleed taps.The engine bleed control system further includes a controller operableto selectively open and close each of the valves based on a bleed airdemand and control delivery of the bleed air to an aircraft use.

According to another embodiment, a gas turbine engine of an aircraft isprovided. The gas turbine engine includes a fan section, a compressorsection, a turbine section, and an engine bleed control system. Theengine bleed control system includes a multi-tap bleed array including aplurality of engine bleed taps coupled to a compressor source of a lowerpressure compressor section before a highest pressure compressor sectionof the gas turbine engine. The highest stage of the engine bleed tapshas a maximum bleed temperature below an auto-ignition point of afuel-air mixture of the aircraft at idle engine power at a maximumaircraft altitude and a pressure suitable for pressurizing the aircraftat the maximum aircraft altitude. The engine bleed control system alsoincludes a plurality of valves operable to extract bleed air from eachof the engine bleed taps. The engine bleed control system furtherincludes a controller operable to selectively open and close each of thevalves based on a bleed air demand and control delivery of the bleed airto an aircraft use.

According to a further embodiment, a method of controlling an enginebleed system for a gas turbine engine of an aircraft is provided. Themethod includes establishing multi-tap bleed array including a pluralityof engine bleed taps coupled to a compressor source of a lower pressurecompressor section before a highest pressure compressor section of thegas turbine engine. A highest stage of the engine bleed taps has amaximum bleed temperature below an auto-ignition point of a fuel-airmixture of the aircraft at idle engine power at a maximum aircraftaltitude and a pressure suitable for pressurizing the aircraft at themaximum aircraft altitude. A plurality of valves operable to extractbleed air from each of the engine bleed taps is configured. Each of thevalves is selectively opened and closed based on a bleed air demand tocontrol delivery of the bleed air to an aircraft use.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter which is regarded as the present disclosure isparticularly pointed out and distinctly claimed in the claims at theconclusion of the specification. The foregoing and other features, andadvantages of the present disclosure are apparent from the followingdetailed description taken in conjunction with the accompanying drawingsin which:

FIG. 1 is a cross-sectional view of a gas turbine engine;

FIG. 2 is a partial view of an engine bleed system according to anembodiment of the disclosure;

FIG. 3 is a schematic view of an aircraft ice control system accordingto an embodiment of the disclosure;

FIG. 4 is a process flow of a method according to embodiments of thedisclosure; and

FIG. 5 is a partial schematic view of another example of a gas turbineengine.

While the above-identified drawing figures set forth one or moreembodiments of the invention, other embodiments are also contemplated.In all cases, this disclosure presents the invention by way ofrepresentation and not limitation. It should be understood that numerousother modifications and embodiments can be devised by those skilled inthe art, which fall within the scope and spirit of the principles of theinvention. The figures may not be drawn to scale, and applications andembodiments of the present disclosure may include features andcomponents not specifically shown in the drawings. Like referencenumerals identify similar structural elements.

DETAILED DESCRIPTION

Various embodiments of the present disclosure are related to enginebleed control for a gas turbine engine. Embodiments of this disclosuremay be applied on any turbomachinery from which compressed air is tappedoff for auxiliary uses. For example, gas turbine engines are rotary-typecombustion turbine engines built around a power core made up of acompressor, combustor and turbine, arranged in flow series with anupstream inlet and downstream exhaust. The compressor compresses airfrom the inlet, which is mixed with fuel in the combustor and ignited togenerate hot combustion gas. The turbine extracts energy from theexpanding combustion gas, and drives the compressor via a common shaft.Energy is delivered in the form of rotational energy in the shaft,reactive thrust from the exhaust, or both. Compressed air can beextracted from various stages as bleed air.

Gas turbine engines provide efficient, reliable power for a wide rangeof applications, including aviation and industrial power generation.Smaller-scale engines such as auxiliary power units typically utilize aone-spool design, with co-rotating compressor and turbine sections.Larger-scale jet engines and industrial gas turbines are generallyarranged into a number of coaxially nested spools, which operate atdifferent pressures and temperatures, and rotate at different speeds.

The individual compressor and turbine sections in each spool aresubdivided into a number of stages, which are formed of alternating rowsof rotor blade and stator vane airfoils. The airfoils are shaped toturn, accelerate and compress the working fluid flow, or to generatelift for conversion to rotational energy in the turbine.

Aviation applications include turbojet, turbofan, turboprop andturboshaft engines. In turbojet engines, thrust is generated primarilyfrom the exhaust. Modern fixed-wing aircraft generally employ turbofanand turboprop designs, in which the low pressure spool is coupled to apropulsion fan or propeller in turbofan with two turbines.Alternatively, in turbofans with three turbines, one turbine drives thefan, one turbine drives the first compressor section and the thirdturbine drives the second compressor section. Turboshaft engines aretypically used on rotary-wing aircraft, including helicopters.

Turbofan engines are commonly divided into high and low bypassconfigurations. High bypass turbofans generate thrust primarily from thefan, which drives airflow through a bypass duct oriented around theengine core. This design is common on commercial aircraft and militarytransports, where noise and fuel efficiency are primary concerns. Lowbypass turbofans generate proportionally more thrust from the exhaustflow, providing greater specific thrust for use on high-performanceaircraft, including supersonic jet fighters. Unducted (open rotor)turbofans and ducted propeller engines are also known, in a variety ofcounter-rotating and aft-mounted configurations.

Referring now to FIG. 1, a cross-sectional view of a gas turbine engine10, in a turbofan configuration is illustrated. The illustrated gasturbine engine 10 includes a fan section 11 with a propulsion fan 12mounted inside a bypass duct 14 upstream of a fan exit guide vane 13. Apower core of the engine is formed by a compressor section 16, acombustor 18 and a turbine section 20. Rotor blades (or airfoils) 21 inthe compressor section 16 and/or the turbine section 20 are arranged instages 38 with corresponding stator vanes (or airfoils) 39, where eachstage 38 includes a rotor blade 21 and stator vane 39 pair.

In the two-spool, high bypass configuration of FIG. 1, compressorsection 16 includes a low pressure compressor 22 (a lower pressurecompressor section) and a high pressure compressor 24 (a highestpressure compressor section). The turbine section 20 includes high apressure turbine 26 and a low pressure turbine 28.

The low pressure compressor 22 is rotationally coupled to the lowpressure turbine 28 via a low pressure shaft 30, thereby forming the lowpressure spool or low spool 31. High pressure compressor 24 isrotationally coupled to the high pressure turbine 26 via a high pressureshaft 32, forming the high pressure spool or high spool 33.

During operation of the gas turbine engine 10, the fan 12 acceleratesair flow from an inlet 34 through bypass duct 14, generating thrust. Thecore airflow is compressed in the low pressure compressor 22 and thehigh pressure compressor 24 and then the compressed airflow is mixedwith fuel in the combustor 18 and ignited to generate combustion gas.

The combustion gas expands to drive the high and low pressure turbines26 and 28, which are rotationally coupled to high pressure compressor 24and low pressure compressor 22, respectively. Expanded combustion gasesexit through exhaust nozzle 36, which is shaped to generate additionalthrust from the exhaust gas flow.

In advanced turbofan designs with a low pressure turbine and a highpressure turbine, the low pressure shaft 30 may be coupled to a lowpressure compressor and then to a fan 12 via geared drive mechanism 37,providing improved fan speed control for increased efficiency andreduced engine noise as a geared turbofan engine. Propulsion fan 12 mayalso function as a first-stage compressor for gas turbine engine 10,with low pressure compressor 22 performing as an intermediate-stagecompressor or booster in front of the high pressure compressor 24.Alternatively, the low pressure compressor stages are absent, and airfrom fan 12 is provided directly to high pressure compressor 24, or toan independently rotating intermediate compressor spool.

An engine accessory gearbox 40 is mechanically coupled via a tower shaft42 to a rotating portion of the gas turbine engine 10, such as the highpressure spool 33. Rotation of various engine accessories, such as pumps44 and electric generators 46 (also referred to as engine generators46), can be driven through the engine accessory gearbox 40 as depictedschematically in FIG. 1. The engine accessory gearbox 40 canalternatively be coupled to low spool 31, and thus the electricgenerators 46 may also be referred to as low spool generators powered byrotation of the low pressure turbine 28 (i.e., lowest pressure turbine).

The gas turbine engine 10 may have a range of different shaft and spoolgeometries, including one-spool, two-spool and three-spoolconfigurations, in both co-rotating and counter-rotating designs. Gasturbine engine 10 may also be configured as a low bypass turbofan, anopen-rotor turbofan, a ducted or un-ducted propeller engine, or anindustrial gas turbine.

FIG. 5 depicts another example of a gas turbine engine 220 in a gearedturbofan configuration. The gas turbine engine 220 extends along anaxial centerline 222 between an upstream airflow inlet 224 and adownstream airflow exhaust 226. The gas turbine engine 220 includes afan section 228, a compressor section 216, a combustor section 232 and aturbine section 219. The compressor section 216 includes a low pressurecompressor (LPC) section 229, an intermediate pressure compressor (IPC)section 230 and a high pressure compressor (HPC) section 231, where theLPC section 229 and IPC section 230 are lower pressure compressorsection before the highest pressure compressor section of HPC section231. The turbine section 219 includes a high pressure turbine (HPT)section 233, an intermediate pressure turbine (IPT) section 234 and alow pressure turbine (LPT) section 235.

The engine sections 228-235 are arranged sequentially along thecenterline 222 within an engine housing 236. This housing 236 includesan inner (e.g., core) casing 238 and an outer (e.g., fan) casing 240.The inner casing 238 houses the LPC section 229 and the engine sections230-235, which form a multi-spool core of the gas turbine engine 220.The outer casing 240 houses at least the fan section 228. The enginehousing 236 also includes an inner (e.g., core) nacelle 242 and an outer(e.g., fan) nacelle 244. The inner nacelle 242 houses and provides anaerodynamic cover for the inner casing 238. The outer nacelle 244 housesand provides an aerodynamic cover the outer casing 240. The outernacelle 244 also overlaps a portion of the inner nacelle 242 therebydefining a bypass gas path 246 radially between the nacelles 242 and244. The bypass gas path 246, of course, may also be partially definedby the outer casing 240 and/or other components of the gas turbineengine 220.

Each of the engine sections 228-231 and 233-235 includes a respectiverotor 248-254. Each of these rotors 248-254 includes a plurality ofrotor blades (e.g., fan blades, compressor blades or turbine blades)arranged circumferentially around and connected to one or morerespective rotor disks. The rotor blades, for example, may be formedintegral with or mechanically fastened, welded, brazed, adhered and/orotherwise attached to the respective rotor disk(s).

The rotors 248-254 are respectively configured into a plurality ofrotating assemblies 256-258. The first rotating assembly 256 includesthe fan rotor 248, the LPC rotor 249 and the LPT rotor 254. The firstrotating assembly 256 can also include a gear train 260 and one or moreshafts 262 and 263, which gear train 260 may be configured as anepicyclic gear train with a planetary or star gear system. The LPC rotor249 is connected to the fan rotor 248. The fan rotor 248 is connected tothe gear train 260 through the fan shaft 262. The LPC rotor 249 istherefore connected to the gear train 260 through the fan rotor 248 andthe fan shaft 262. The gear train 260 is connected to and driven by theLPT rotor 254 through the low speed shaft 263.

The second rotating assembly 257 includes the IPC rotor 250 and the IPTrotor 253. The second rotating assembly 257 also includes anintermediate speed shaft 264. The IPC rotor 250 is connected to anddriven by the IPT rotor 253 through the intermediate speed shaft 264.

The third rotating assembly 258 includes the HPC rotor 251 and the HPTrotor 252. The third rotating assembly 258 also includes a high speedshaft 265. The HPC rotor 251 is connected to and driven by the HPT rotor252 through the high speed shaft 265.

One or more of the shafts 262-265 may be coaxial about the centerline222. One or more of the shafts 263-265 may also be concentricallyarranged. The low speed shaft 263 is disposed radially within andextends axially through the intermediate speed shaft 264. Theintermediate speed shaft 264 is disposed radially within and extendsaxially through the high speed shaft 265. The shafts 262-265 arerotatably supported by a plurality of bearings; e.g., rolling elementand/or thrust bearings. Each of these bearings is connected to theengine housing 236 (e.g., the inner casing 238) by at least onestationary structure such as, for example, an annular support strut.

During operation, air enters the gas turbine engine 220 through theairflow inlet 224. This air is directed through the fan section 228 andinto a core gas path 266 and the bypass gas path 246. The core gas path266 flows sequentially through the engine sections 229-235. The airwithin the core gas path 266 may be referred to as “core air”. The airwithin the bypass gas path 246 may be referred to as “bypass air”.

The core air is compressed by the compressor rotors 249-251 and directedinto the combustor section 232. Fuel is injected into the combustorsection 232 and mixed with the compressed core air to provide a fuel-airmixture. This fuel air mixture is ignited and combustion productsthereof flow through and sequentially cause the turbine rotors 252-254to rotate. The rotation of the turbine rotors 252-254 respectively driverotation of the compressor rotors 251-249 and, thus, compression of theair received from a core airflow inlet. The rotation of the turbinerotor 254 also drives rotation of the fan rotor 248, which propelsbypass air through and out of the bypass gas path 246. The propulsion ofthe bypass air may account for a majority of thrust generated by the gasturbine engine 220, e.g., more than seventy-five percent (75%) of enginethrust. The gas turbine engine 220 of the present disclosure, however,is not limited to the foregoing exemplary thrust ratio. Further,although the example of FIG. 5 includes gear train 260, the gear train260 can be eliminated in other embodiments that include two or morespools.

FIG. 2 is a partial view of an engine bleed system 50 (also referred toas an engine bleed control system) according to an embodiment. In theexample of FIG. 2, the engine bleed system 50 includes a multi-tap bleedarray 51 which includes a plurality of engine bleed taps 52A, 52B, 52C,52D coupled to a compressor source 54 of the gas turbine engine 10. Eachof the engine bleed taps 52A-52D can be located at a lower pressurelocation, for instance, before a ninth stage 55 of rotor blade 21 andstator vane 39 pairs of a compressor section 16 of the gas turbineengine 10. In some embodiments, the compressor source 54 is the lowestpressure compressor source of compressor section 16. Although theexample of FIG. 2 depicts four engine bleed taps 52A-52D, it will beunderstood that the multi-tap bleed array 51 can include any number oftwo or more engine bleed taps between a fan-air source 56 and a highestpressure compressor section of compressor section 16 in variousembodiments. Engine bleed tap 52A is at an upstream location withrespect to engine bleed taps 52B-52D and thus provides a source of lowercompression and cooler bleed air as compared to bleed air extracted fromengine bleed taps 52B-52D. Similarly, engine bleed tap 52B is at anupstream location with respect to engine bleed taps 52C-52D and thusprovides a source of lower compression and cooler bleed air as comparedto bleed air extracted from engine bleed taps 52C-52D. Engine bleed tap52D is at a downstream location with respect to engine bleed taps52A-52C and thus provides a source of higher compression and hotterbleed air as compared to bleed air extracted from engine bleed taps52A-52C. In embodiments, a highest stage of the engine bleed taps (i.e.,engine bleed tap 52D) has a maximum bleed temperature below anauto-ignition point of a fuel-air mixture of the aircraft at idle enginepower at a maximum aircraft altitude and a pressure suitable forpressurizing the aircraft at the maximum aircraft altitude. A loweststage of the engine bleed taps (i.e., engine bleed tap 52A) has amaximum bleed temperature below the auto-ignition point of the fuel-airmixture of the aircraft at a highest engine power operation and apressure suitable for pressurizing the aircraft.

In the example of FIG. 2, bleed air from engine bleed tap 52A is routedthrough a check valve 58A to intermediate duct 59. A valve 62A cancontrol delivery of the bleed air from intermediate duct 59 to anaircraft use 64 through ducts 65. Valve 62A can be a shutoff valve or acombined pressure regulating and shutoff valve. The aircraft use 64 maybe an environmental control system 90 of an aircraft 5, as best seen inFIG. 3. Bleed air from engine bleed tap 52B can be routed through checkvalve 58B to intermediate duct 59 as controlled by valve 62B. Bleed airfrom engine bleed tap 52C can be routed through check valve 58C tointermediate duct 59 as controlled by valves 62B and 62C. Bleed air fromengine bleed tap 52D can be routed to intermediate duct 59 as controlledby valves 62B, 62C, and 62D. Other configurations of the multi-tap bleedarray 51 are contemplated, including different valve arrangements with agreater or lesser number of valves. For example, rather than cascadingvalves 62B-62D, valve 62C and/or valve 62D can be directly connected tointermediate duct 59.

In embodiments, a pneumatic bleed 70 for anti-icing a nacelle inlet 72(FIG. 3) of the gas turbine engine 10 is provided for an engineanti-icing system 74. The engine anti-icing system 74 can provideanti-icing for engine components and/or nacelle components and canexceed 400 degrees Fahrenheit (204 degrees Celsius). The pneumatic bleed70 can be at a different engine stage than the engine bleed taps52A-52D, e.g., higher temperature/compression point downstream, but neednot be located at the highest stage of compression. A valve 76 can beselectively actuated by a controller 48 to enable the engine anti-icingsystem 74. In some embodiments, a wing anti-icing system 78 in wing 80of the aircraft 5 is powered by an engine generator 46, i.e., electricanti-icing. In alternate embodiments, the controller 48 is operable tocontrol delivery of a portion of the bleed air to the wing anti-icingsystem 78 of the aircraft 5 using valve 82. The controller 48 may alsocontrol valves 62A-62D, as well as other components.

The controller 48 may include memory to store instructions that areexecuted by a processor. The executable instructions may be stored ororganized in any manner and at any level of abstraction, such as inconnection with a controlling and/or monitoring operation of one or moresystems of the gas turbine engine 10 of FIG. 1. The processor can be anytype of central processing unit (CPU), including a general purposeprocessor, a digital signal processor, a microcontroller, an applicationspecific integrated circuit (ASIC), a field programmable gate array, orthe like. Also, in embodiments, the memory may include random accessmemory (RAM), read only memory (ROM), or other electronic, optical,magnetic, or any other computer readable medium onto which is storeddata and control algorithms in a non-transitory form. The controller 48can be embodied in an individual line-replaceable unit, within a controlsystem (e.g., in an electronic engine control), and/or distributedbetween multiple electronic systems.

In the example of FIG. 2, source locations of the engine bleed taps52A-52D are selected to hold a maximum temperature of the bleed airbelow an auto-ignition point of a fuel-air mixture at all flightconditions of the gas turbine engine 10. For instance, the maximumtemperature can be established as 400 degrees Fahrenheit (204 degreesCelsius) for 0.25 mach and a 120 degree Fahrenheit day. The controller48 may observe various aircraft operating conditions to determinepressures and temperatures at each of the engine bleed taps 52A-52D andselectively open and close valves 62A-62D based on a bleed air demandand control delivery of the bleed air to aircraft use 64 and/or winganti-icing system 78.

While a specific configuration is depicted in FIG. 2, otherconfigurations are contemplated within the scope of embodiments. Forinstance, the valve 82 may be located upstream of one or more of valves62A-62D. Further, output of one or more of the engine bleed taps 52A-52Dmay have other uses and/or connections with the wing anti-ice system 78and/or other systems. The multi-tap bleed array 51 and/or valves 62A-62Dmay be located proximate to the gas turbine engine 10, below or within apylon 84 (FIG. 3) that couples a nacelle of the gas turbine engine 10 towing 80, or within the aircraft 5. Further, the engine bleed system 50can be incorporated into the gas turbine engine 220 of FIG. 5, whereengine bleed tap 52 can be coupled to a compressor source of a lowerpressure compressor section (e.g., LPC section 229 or IPC section 230)before a highest pressure compressor section (HPC section 231) of thegas turbine engine 220 of FIG. 5, for example.

FIG. 4 is a process flow of a method 100 according to an embodiment. Themethod 100 is described with reference to FIGS. 1-5. Although describedprimarily in reference to the gas turbine engine 10 of FIG. 1, it willbe understood that the method 100 can also be applied to the gas turbineengine 220 of FIG. 5 and other configurations. At block 102, multi-tapbleed array 51 is established with a plurality of engine bleed taps52A-52D coupled to a compressor source 54 of a lower pressure compressorsection before a highest pressure compressor section of the gas turbineengine 10, where a highest stage of the engine bleed taps (e.g., enginebleed tap 52D) has a maximum bleed temperature below an auto-ignitionpoint of a fuel-air mixture of the aircraft at idle engine power at amaximum aircraft altitude and a pressure suitable for pressurizing theaircraft at the maximum aircraft altitude. At block 104, valves 62A-62Dare configured to extract bleed air from each of the engine bleed taps52A-52D. Check valves 58A-58C can also control the flow of bleed airfrom engine bleed taps 52A-52D to intermediate duct 59, for instance. Atblock 106, each of the valves 62A-62D is selectively opened and closedbased on a bleed air demand to control delivery of the bleed air to anaircraft use 64. Anti-icing can be provided from a pneumatic bleed 70 toa nacelle inlet 74 of the gas turbine engine 10. Power from an enginegenerator 46 can be provided to a wing anti-icing system 78 of theaircraft 10. Alternatively, controller 48 controls delivery of a portionof the bleed air to the wing anti-icing system 78 of the aircraft 5,e.g., using a combination of valves 62A-62D and/or 82.

Technical effects and benefits include reducing engine bleed energy lossusing multiple engine bleed taps and a peak temperature limit.Embodiments selectively open and close valves based on a bleed airdemand to maintain pressure and temperature limits and avoid precoolingthe engine bleed air. Embodiments can eliminate the need for apre-cooler or additional heat exchanger by selecting engine bleed airfrom an engine tap at targeted temperature and pressure while notexceeding the auto-ignition point of a fuel-air mixture.

While the present disclosure has been described in detail in connectionwith only a limited number of embodiments, it should be readilyunderstood that the present disclosure is not limited to such disclosedembodiments. Rather, the present disclosure can be modified toincorporate any number of variations, alterations, substitutions orequivalent arrangements not heretofore described, but which arecommensurate with the scope of the present disclosure. Additionally,while various embodiments of the present disclosure have been described,it is to be understood that aspects of the present disclosure mayinclude only some of the described embodiments. Accordingly, the presentdisclosure is not to be seen as limited by the foregoing description,but is only limited by the scope of the appended claims.

The invention claimed is:
 1. A method of controlling an engine bleedsystem for a gas turbine engine of an aircraft, the method comprising:establishing a multi-tap bleed array comprising a plurality of enginebleed taps distributed axially between a plurality of stages of acompressor section of the gas turbine engine, wherein each stage of theplurality of stages comprises a rotor blade and stator vane pair, theplurality of engine bleed taps comprises at least four bleed taps, andthe plurality of engine bleed taps is located before a ninth stage ofthe plurality of stages of the compressor section; configuring aplurality of valves operable to extract bleed air from each of theplurality of engine bleed taps; selectively opening and closing each ofthe plurality of valves based on a bleed air demand to control deliveryof the bleed air to an aircraft use and in response to a controller; andproviding, responsive to the controller, from a pneumatic bleed of thecompressor section to at least a nacelle inlet of the gas turbineengine, wherein the pneumatic bleed is located downstream at a highertemperature point than the plurality of engine bleed taps.
 2. The methodas in claim 1, wherein the aircraft use is an environmental controlsystem of the aircraft that receives the bleed air as controlled by thecontroller and the plurality of engine bleed taps has a maximum bleedtemperature below 400 degrees Fahrenheit (204 degrees Celsius) at ahighest engine power operation and a pressure suitable for pressurizingthe aircraft.
 3. The method as in claim 1, further comprising:controlling, using the controller, a valve that controls delivery of aportion of the bleed air to an anti-icing system of the aircraft.
 4. Anengine bleed control system for an aircraft, the engine bleed controlsystem comprising: a gas turbine engine comprising a compressor section;a multi-tap bleed array comprising a plurality of engine bleed tapsdistributed axially between a plurality of stages of the compressorsection, wherein each stage of the plurality of stages comprises a rotorblade and stator vane pair, the plurality of engine bleed taps comprisesat least four bleed taps, and the plurality of engine bleed taps islocated before a ninth stage of the plurality of stages of thecompressor section; a plurality of valves configured to extract bleedair from each of the plurality of engine bleed taps; a controllercomprising a processor and memory programmed with a plurality ofinstructions to selectively open and close each of the plurality ofvalves based on a bleed air demand and control delivery of the bleed airto an aircraft use; and a pneumatic bleed of the compressor sectionlocated downstream at a higher temperature point than the plurality ofengine bleed taps, wherein the pneumatic bleed provides anti-icing to atleast a nacelle inlet of the gas turbine engine responsive to executionof the plurality of instructions by the processor of the controller. 5.The engine bleed control system as in claim 4, wherein the aircraft useis an environmental control system of the aircraft that receives thebleed air as controlled by the controller.
 6. The engine bleed controlsystem as in claim 4, further comprising an engine generator controlledby the controller, wherein the engine generator is configured to power awing anti-icing system of the aircraft.
 7. The engine bleed controlsystem as in claim 4, wherein the controller is configured to control avalve that controls delivery of a portion of the bleed air to ananti-icing system of the aircraft.
 8. The engine bleed control system asin claim 4, wherein the plurality of engine bleed taps has a maximumbleed temperature below 400 degrees Fahrenheit (204 degrees Celsius) ata highest engine power operation and a pressure suitable forpressurizing the aircraft.
 9. The engine bleed control system as inclaim 4, wherein the multi-tap bleed array and the plurality of valvesare configured to be located below a pylon coupling a nacelle of the gasturbine engine to a wing of the aircraft.
 10. The engine bleed controlsystem as in claim 4, wherein the gas turbine engine is a gearedturbofan engine, and the gas turbine engine comprises a low-turbinepowered electric generator configured to power a wing anti-icing systemof the aircraft.
 11. The engine bleed control system of claim 4, furthercomprising an intermediate duct, a shutoff valve coupled to the aircraftuse, a second shutoff valve coupled to the intermediate duct, a thirdshutoff valve coupled to the second shutoff valve, and a fourth shutoffvalve coupled to the third shutoff valve, wherein a first engine bleedtap of the plurality of engine bleed taps is routed through a firstcheck valve to the intermediate duct, a second bleed tap of theplurality of engine bleed taps is routed through a second check valve tothe second shutoff valve, a third bleed tap of the plurality of enginebleed taps is routed through a third check valve to the third shutoffvalve, and a fourth bleed tap of the plurality of engine bleed taps isrouted to the fourth shutoff valve.
 12. A gas turbine engine for anaircraft, the gas turbine engine comprising: a fan section; a compressorsection; a turbine section; and an engine bleed control systemcomprising: a multi-tap bleed array comprising a plurality of enginebleed taps distributed axially between a plurality of stages of thecompressor section, wherein each stage of the plurality of stagescomprises a rotor blade and stator vane pair, the plurality of enginebleed taps comprises at least four bleed taps, and the plurality ofengine bleed taps is located before a ninth stage of the plurality ofstages of the compressor section; a plurality of valves configured toextract bleed air from each of the plurality of engine bleed taps; acontroller comprising a processor and memory programmed with a pluralityof instructions to selectively open and close each of the plurality ofvalves based on a bleed air demand and control delivery of the bleed airto an aircraft use; and a pneumatic bleed of the compressor sectionlocated downstream at a higher temperature point than the plurality ofengine bleed taps, wherein the pneumatic bleed provides anti-icing to atleast a nacelle inlet of the gas turbine engine responsive to executionof the plurality of instructions by the processor of the controller. 13.The gas turbine engine as in claim 12, wherein the aircraft use is anenvironmental control system of the aircraft that receives the bleed airas controlled by the controller.
 14. The gas turbine engine as in claim12, further comprising an engine generator controlled by the controller,wherein the engine generator is configured to power a wing anti-icingsystem of the aircraft.
 15. The gas turbine engine as in claim 12,wherein the controller is configured to control a valve that controlsdelivery of a portion of the bleed air to an anti-icing system of theaircraft.
 16. The gas turbine engine as in claim 12, wherein theplurality of engine bleed taps has a maximum bleed temperature below 400degrees Fahrenheit (204 degrees Celsius) at a highest engine poweroperation and a pressure suitable for pressurizing the aircraft.
 17. Thegas turbine engine as in claim 12, wherein the multi-tap bleed array andthe plurality of valves are configured to be located below a pyloncoupling a nacelle of the gas turbine engine to a wing of the aircraft,the gas turbine engine is a geared turbofan engine, and the gas turbineengine comprises a low-turbine powered electric generator configured topower a wing anti-icing system of the aircraft.
 18. The gas turbineengine of claim 12, further comprising an intermediate duct, a shutoffvalve coupled to the aircraft use, a second shutoff valve coupled to theintermediate duct, a third shutoff valve coupled to the second shutoffvalve, and a fourth shutoff valve coupled to the third shutoff valve,wherein a first engine bleed tap of the plurality of engine bleed tapsis routed through a first check valve to the intermediate duct, a secondbleed tap of the plurality of engine bleed taps is routed through asecond check valve to the second shutoff valve, a third bleed tap of theplurality of engine bleed taps is routed through a third check valve tothe third shutoff valve, and a fourth bleed tap of the plurality ofengine bleed taps is routed to the fourth shutoff valve.